Double-wall airfoil

ABSTRACT

A double-wall airfoil for applications such as the blades and vanes of gas turbine engines. The double-wall comprises an outer airfoil skin and an inner support wall that are metallurgically bonded to one another. The double-wall contains integral channels for passage of cooling air adjacent to the airfoil skin. Airfoil skin may be a metal alloy skin or a microlaminate structure, including microlaminate composite structures. Microlaminate composites typically have a lower density than that of the material used for the airfoil support wall, and a simplified internal geometry which promote weight reductions in the airfoils and increases in engine operating efficiency.

FIELD OF THE INVENTION

The present invention relates generally to airfoils, such as turbine orcompressor blades or vanes, for use at high temperatures. Moreparticularly, this invention pertains to airfoils having a double-wallconstruction with integral cooling channels. Most particularly, thisinvention pertains to double-wall airfoils having a micro-laminate outerwall that is adapted to withstand high temperatures, such as are foundin the hot sections of gas turbines and aircraft engines.

BACKGROUND OF THE INVENTION

It is known that the temperatures of combustion gases in gas turbineengines, such as aircraft engines, during operation are considerablyabove the temperatures of the metal parts of the engine which are incontact with these gases. Operation of these engines at gas temperaturesthat are above the metal part temperatures is a well established art,and depends on supplying cooling gas to the outer surfaces of the metalparts through various methods. The metal parts of these engines that areparticularly subject to high temperatures, and thus require particularattention with respect to cooling, are the blades and vanes used todirect the flow of the hot gases

For example, with regard to the metal blades and vanes employed inaircraft engines, some cooling is achieved by convection by providingpassages for flow of a cooling gas internally within the blades so thatheat may be removed from the metal structure of the blade by the coolinggases. Such blades are essentially hollow core blades which may have ashell or plural shells of structural metal of intricate designsurrounding equally intricate sets of cooling passages within the hollowcore blade. Fine internal orifices have also been devised to direct thiscirculating cooling gas directly against certain inner surfaces of theshell to obtain cooling of the inner surface by impingement of thecooling gas against the surface, a process known as impingement cooling.In addition, an array of fine holes extending from the hollow corethrough the blade shell can provide for bleeding cooling air through theblade shell to the outer surface where a film of such air can protectthe blade from direct contact with the hot gases passing through theengines, a process known as film cooling.

Using combinations of these cooling techniques, the maximum metalsurface temperature of a blade can be maintained at about 1,150° C.while the blade is in an environment of hot gases having temperatures ofup to 1,650° C.

As is well-known, the operating efficiency of gas turbines andaccordingly of jet engines is related to the operating temperature ofthe engine. To achieve higher operating efficiencies, operation of theengine at higher temperatures is desirable. For engines operating attemperatures up to 2,000° C., it is expected that the metal temperaturescan be maintained almost at present levels with current coolingtechniques by using a combination of improved cooling designs andthermal barrier coatings. Thermal barrier coatings are well-knownceramic coatings such as yttria stabilized zirconia that are applied tothe external surface of metal parts within engines to impede thetransfer of heat from hot combustion gases to the metal parts. However,even with the use of advanced gas cooling designs and thermal barriercoatings, it is also desirable to decrease the requirement for coolinggases, because reducing the demand for cooling gases is also well-knownto improve overall engine operating efficiency. One way to achieve sucha reduction is to improve the cooling of the metal parts immediatelyadjacent to their outer surfaces.

Another way in which the increased use of cooling air can be avoided, orcooling air requirements can be reduced, is by providing metal partsthat are capable of operating above the maximum use temperature of1,150° C. The provision of metal parts capable of operating attemperatures beyond 1,150° C. would allow either relaxation of coolingrequirement or the reduction or elimination of the dependence on thethermal barrier coatings, or both.

It is also well-known that the operating efficiency of gas turbineengines may be improved by reducing the total weight of the metal partsutilized. Currently, because of the required intricate internal coolingpassages within metal parts such as blades and vanes, particularly neartheir outer surfaces, and the fragile nature of the ceramic cores usedto define these passages during formation, it is necessary to utilizelarge tolerances that allow for the possibility of core shifting. Theuse of materials and processes that would simplify the designrequirements for these internal passages would permit the amount ofmaterial used in each metal part to be reduced. Also, the use ofmaterials that are less dense would achieve weight reductions for eachmetal part. Small savings can be significant because of the large numberof these metal parts that are utilized in a typical engine.

Reducing the internal complexity of the metal parts by reducing thenumber of intricate passageways that must be formed by casting wouldalso improve casting yields and provide an added benefit.

Therefore, it is desirable to define airfoils and materials for theirmanufacture that have improved cooling capability, higher operatingtemperatures, more castable geometries and reduced weight as compared topresent airfoils.

SUMMARY OF THE INVENTION

The present invention is a double-wall airfoil having at least oneintegral, longitudinally-extending cooling channel located in thedouble-wall. The integral cooling channel comprises a means to provideadditional cooling capacity to the exterior of the airfoil. An airfoilskin is deposited on a airfoil support wall to define the integralchannels. It is preferred that the airfoil skin be a microlaminatecomposite of a more ductile phase and a less ductile phase, where theless ductile phase also has significant high temperature strength. Themore ductile phase may comprise an intermetallic compound orintermediate phase having significant high temperature strength.

This invention comprises an airfoil having an outer double-wall,comprising: a partially-hollow airfoil support wall that is attached toand extends longitudinally from an airfoil base and has anairfoil-shaped outer surface, said airfoil support wall formed from afirst material and having at least one longitudinally-extending recessedchannel formed in the outer surface; and an airfoil skin made from asecond material which conforms to, covers and is metallurgically bondedto the airfoil-shaped outer surface of said airfoil support wall andcovers the at least one recessed channel, wherein the combination ofsaid airfoil skin and said airfoil support wall form a double-wallairfoil structure and the covered, recessed channel forms at least oneintegral internal channel located within the double-wall.

A significant advantage of this invention is that it provides airfoilsthat may be utilized in hotter operating environments than are possiblewith prior art airfoils, thereby permitting a turbine engine thatincorporates them to potentially be operated at temperatures thatprovide greater engine operating efficiencies.

Another significant advantage is that airfoils of the present inventionmay potentially be lighter due to for example, decreased densities ofmaterials used for the airfoil skin and material reductions related tosimpler internal geometries and smaller manufacturing tolerances.

Other objects, features and advantages of airfoils of this invention maybe seen in the description and drawings set forth below, including theappended claims.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic cross-sectional illustration of a portion of a gasturbine engine illustrating a representative arrangement of airfoils ofthe present invention as a blade and vane.

FIG. 1A is schematic cross-sectional illustration of an airfoil of thepresent invention.

FIG. 2 is a schematic cross-sectional illustration of a secondembodiment of an airfoil of the present invention, representing apotential turbine blade configuration.

FIG. 3 is a schematic cross-sectional illustration of a third embodimentof an airfoil of the present invention, representing a potential turbinevane airfoil configuration.

FIG. 4 is a plot in which the projected yield strength and 100 hr/1%creep limits as a function of temperature generally comparing presentairfoil alloys with intermetallic composites.

FIG. 5 is a schematic diagram which illustrates a method of makingdouble-wall airfoils.

DETAILED DESCRIPTION OF THE INVENTION

The present invention relates generally to airfoils, such as turbineblades or vanes, for use at high temperatures. More particularly, thisinvention pertains to airfoils having a double-wall construction with alongitudinally-extending integral cooling channel or channels. Mostparticularly, this invention pertains to double-wall airfoils having amicro-laminate outer wall that is adapted to withstand extremely hightemperatures, such as are found in the hot sections of aircraft enginesand other gas turbines.

Airfoil Structure

The airfoils of this invention are suitable for use in either therotating or stationary portions of a turbine engine. They are preferablyused in the hot sections of such engines where they are exposed tocombustion gases having a temperature of about 1650° C. or more.

Referring now to FIGS. 1 and 1A, a portion of a turbine engine 10 isshown illustrating a blade 12 and vane 14 of the present invention.Blade 12 and vane 14 are of known external configuration, however,airfoils of the present invention may also be utilized with differentblade and vane designs, including advanced designs. Blade 12 comprisesairfoil 16 that is attached to base 18, as explained further herein.Vane 14 comprises airfoil 20 that is attached to base 22, as explainedfurther herein. Airfoils 16 and 20 (subsequent references are to airfoil16 comprise the same essential internal elements, and thus are describedwith reference to the same FIG. 1A. Being of known externalconfiguration, prior art airfoils and airfoil 16,20 of this inventionmay be described generally as comprising a structural member, frequentlyof complex shape, that extends generally longitudinally from a base 18or base 22 (subsequent references are to base 18, but apply to either abase for a blade or a vane) and has a generally convex leading edge 24connected on one side to a generally concave pressure side wall 26 andon the other side to generally convex suction side wall 28, saidpressure side wall 26 and said suction side wall 28 converging to formtrailing edge 30. The combination of these elements define anairfoil-shape as utilized herein. Further, airfoils 16, 20 comprise apartially-hollow airfoil support wall 40 and airfoil skin 42. Airfoilsupport wall 40 extends longitudinally in direction 44 from base 18.Airfoil support wall 40 is made from a first material and has at leastone longitudnally-extending recessed channel 46 formed in anairfoil-shaped outer surface 48. Airfoil skin 42 conforms toairfoil-shaped outer surface 48 of airfoil support wall 40. Airfoil skin42 covers the at least one channel 46 in airfoil support wall 40, and ismetallurgically bonded to airfoil-shaped outer surface 48 of airfoilsupport wall 40. The combination of airfoil support wall 40 and airfoilskin 42 forms double-wall structure 50. Where recessed channel 46 iscovered by airfoil skin 42, it forms at least one longitudinallyextending internal channel 52. Even though the essential elements ofblade 12 and vane 14 are the same, it will be known to those of ordinaryskill to alter the arrangement of these elements to accommodate theirdiffering design criteria, such as altering the thickness andarrangement of the airfoil support wall, the degree to which the airfoilsupport wail is hollow, the number and placement of the cooling channelsand other design parameters affecting cooling by impingement, convectionor cooling holes in the airfoil skin to form surface films of air.

Airfoil support wall 40 is attached to base 18. Base 18 may be of anysuitable configuration and material that permit attachment to andoperation within turbine engine 10. Bases 18 having certain dovetail orother configurations for interlocking with other components of an engine10 made from Ni-base alloys are known. Preferably, airfoil support wall40 is formed integrally with base 18, such as by casting base 18 andairfoil support wall 40 together as a single part from a singlematerial. Applicants believe that it will be further preferred that base18 and airfoil support wall 40 be cast so as to form a directionallysolidified or single structure. Other forms of attachment may include,without limitation, casting airfoil support wall 40 into a base that isadapted to receive airfoil support wall 40, or attaching a preformedairfoil support wall 40 to a preformed base 18. It is preferred,however, regardless of the means of attachment, that certain knownmechanical properties at the point of attachment, particularlymechanical strength and creep resistance be maintained.

Airfoil support wall 40 extends longitudinally in direction 44 away frombase 18, providing the airfoil with a length. Airfoil support wall 40 isalso partially-hollow as represented by hollow regions or cavities 54.

Hollow regions 54 as well as integral internal channels 52 may be usedin the operation of engine 10 to circulate a cooling gas (not shown),such as air, within airfoil 16 in order to cool these articles. Thecooling gas is typically supplied through orifices in base 18. Incontrast to prior art airfoil designs, hollow regions 54 may be ofrelatively simple designs with cross-sectional areas that avoid thenecessity of forming intricate hollow regions near the surface ofairfoil 16. This is due to the fact that channels 52 are capable ofproviding such intricate passageways, without the need for casting thementirely within airfoil support wall 40. Examples to illustrate thispoint are shown as FIGS. 2 and 3. FIG. 2 illustrates an alternateairfoil support wall 40 cross-section, such as could be used in aturbine blade design. FIG. 3 illustrates an alternate airfoil supportwall 40 cross-section, such as could be used in a turbine vane design

In the case of airfoil 16, it is believed that the cross-section ofairfoil support wall 40 as shown in FIG. 1A will be similar along itslength with regard to its general configuration (e.g. the number ofchannel 52, the number and arrangement of hollow regions 54, and across-sectional area of airfoil 16). However, this is not an essentialfeature of the invention and this invention is intended to comprehendthose designs for which the cross-sectional configuration may vary overthe length of airfoil 16.

Airfoil support wall 40 may be formed from any suitable material,described herein as a first material. Depending on the intendedapplication for airfoil 16, this could include Ni-base, Co-base andFe-base superalloys. Fe-base and Co-base alloys may be preferred forrelatively larger scale engines, such as industrial gas turbines.Ni-base superalloys may be preferred for relatively smaller-scaleapplications, such as aircraft engines, where the operating gastemperatures are up to about 1,650° C. For such applications, thepreferred Ni-base superalloys may be those containing both γ and γ'phases, particularly those Ni-base superalloys containing both γ and γ'phases wherein the γ' phase occupies at least 40% by volume of thesuperalloy. Such alloys are known to be advantageous because of acombination of desirable properties including high temperature strengthand high temperature creep resistance. First material may also comprisea NiAl intermetallic alloy, as these alloys are also known to possess acombination of superior properties including high temperature strengthand high temperature creep resistance that are advantageous for use inturbine engine applications used for aircraft. In the case of theNi-base alloys described, it is preferred that the tensile strength ofthe first material be at least 175 MPa at temperatures in the range of1,100°-1,200° C. For advanced applications (e.g. applications where theoperating temperatures are above about 1,650° C.), it will be preferredto use higher temperature first materials, such as Nb-base alloys. Inthe case of Nb-base alloys, coated Nb-base alloys having superioroxidation resistance will be preferred, such as Nb/Ti alloys, andparticularly those alloys comprisingNb-(27-40)Ti-(4.5-10.5)Al-(4.5-7.9)Cr-(1.5-5.5) Hf-(0-6)V in a atompercent. First material may also comprise a Nb-base alloy that containsat least one secondary phase, such as an Nb-containing intermetalliccompound, an Nb-containing carbide or an Nb-containing boride. Suchalloys are analogous to a composite material in that they contain aductile phase (i.e. the Nb-base alloy) and a strengthening phase (i.e.an Nb-containing intermetallic compound, an Nb-containing carbide or anNb-containing boride).

Referring to FIG. 2, an airfoil support wall 40 of the present inventionwill have at least one longitudinally-extending recessed channel 46formed in airfoil-shaped outer surface 48. It is preferred, however,that airfoil support wall 40 comprise a plurality of recessed channels46 so that a plurality of integral channels 52 can be formed just underthe surface of airfoil skin 42, such that a cooling gas (not shown) maybe circulated through internal channel 52. Recessed channel may beformed by any suitable forming method. It is preferred that recessedchannels 46 be formed integrally with airfoil support wall 40, such asby casting them into airfoil support wall 40. However, recessed channels46 may be formed by other methods that are not integral with theformation of airfoil support wall 40, such as conventional machining,water-jet machining, etching or other suitable methods. Recessedchannels 46 may be of any suitable cross-section. The cross-section ofrecessed channel 46 will depend upon factors including the materialselected as airfoil skin 42 and the thickness of airfoil skin 42. It isnecessary, however, that cross-section of recessed channels 46 bedesigned such that the combination of the cross-section and airfoil skin42 does not permit airfoil skin 42 to be forced into recessed channels46 during use of airfoil 16. It is believed that recessed channels 46that are approximately 0.1-0.3 inches wide and 20-50 mils deep willprovide an adequate channel to permit the flow of a cooling gas whileavoiding the problems described above. It is preferred that recessedchannels 46 occupy a substantial portion of the area of airfoil-shapedouter surface 48, such as about 50 percent or more of the double wallportion of the airfoil skin 42. Recessed channel or channels 46 may haveany suitable configuration over airfoil-shaped outer surface 48,however, it is preferred that recessed channel or channels 46 extendlongitudinally along the length of blade 12 or vane 14 in the generaldirection indicated by arrow 44 as shown in FIG. 1. Further, the atleast one longitudinally extending channel or channels 46 may form manypatterns over airfoil-shaped outer surface 48, such as a serpentinepattern shown in FIG. 1, wherein one continuous channel may extend overa significant portion of airfoil-shaped outer surface 48. By orientatingthe at least one channel 46 longitudinally, a channel 52 or channels canbe formed that run under the length of particularly hot sections, suchas the lengths of leading and trailing edges.

With regard to forming airfoils 16 of the present invention, oneadvantage of the present invention is the simplified geometric form ofthe casting required to form airfoil support wall 40 as compared to themore complex castings required to cast integral double-wall airfoils asdescribed in U.S. Pat. No. 5,328,331, which incorporates intricateinternal cooling passages near the surface of the airfoil. Thissimplified geometric form will permit a substantial increase in themanufacturing yield of airfoil support wall 40 as compared to prior artairfoils when using manufacturing processes such as casting anddirectional solidification. Another advantage related to the simplifiedgeometric form is that it increases the range of superalloys that may beutilized to form airfoil support wall 40 as compared with an integrallyformed double-wall airfoil. Particularly, permits the use of strongersuperalloys which are typically less castable with respect to smallinterior features and thin airfoil walls. In alloy development ofadvanced single crystal and directionally solidified alloys, castabilityis a barrier that eliminates from consideration many very strongmaterials; the simpler cast geometric forms of the present inventionthat do not require the casting of an integral outer surface may permitthe use of some of these stronger alloys. Also, another advantageassociated with simplified geometries is that high casting yields areexpected.

Moreover, because of the novel design and construction of the airfoil ofthis invention as described briefly above, current Ni-base superalloyswhich have higher densities due to the presence of higher concentrationsof refractory metal strengthening additives such as tantalum, which areneeded to achieve strength at the highest temperatures, may be replacedby alloys having lower density strengthening additives, such astitanium. For example, because airfoil support wall 40 in airfoils ofthe present invention may in some cases be exposed to lower temperaturesthan those experienced by prior art airfoils due to the presence ofintegral cooling channels 52, it may be possible to use Ti to stabilizethe strengthening γ phase, rather than a refractory metals such astungsten, tantalum or niobium. The use of lighter alloying elementswould reduce the alloy density, and translate to reduced airfoil weightand higher engine operating efficiency. Also Ti is known to be able tostabilize greater quantities of γ' than can be stabilized by refractorymetals, such that stronger alloys may be utilized.

Another advantage related to the simplified geometry of airfoil supportwall 40 is that it is not dependent upon ceramic core limitations. Inthe manufacture of hollow core airfoils, one of the limitations whichhas been recognized in the industry is that both the cavity geometry andthe cavity sizes are dictated by ceramic core limitations. Since thesacrificial ceramic core used in casting is a brittle structure, itsgeometry and size must be selected so that it withstands the castingprocess, without breaking and without substantial distortion. Anydistortion of the ceramic core must be accounted for in designingtolerances of the cast airfoil wall thickness, including the possibilitythat fine core elements are frequently subject to shifting. Therefore,the designs of the present invention will take advantage of tighterachievable tolerances which translate to reductions in the amount ofmaterial required for a given casting (e.g. internal wall thicknesseswithin the structure may be thinner because they do not have to allowfor the possibility that the core used to form the cavity may shift inplace during casting).

Airfoil skin 42 extends longitudinally along airfoil-shaped outersurface 48 of airfoil support wall 40. Airfoil skin 42 conforms toairfoil-shaped outer surface 48 and covers recessed channels 46. Airfoilskin 42 comprises a second material which may be any suitable materialand is metallurgically bonded to the airfoil-shaped outer surface 48 ofairfoil support wall 40. It is preferred that the thickness of airfoilskin 42 be in the range of 0.010-0.040 inches. However, otherthicknesses may be utilized depending on the requirements for aparticular airfoil.

It is preferred that the airfoil support wall 40 be designed, includingthe selection of the first material, so as to serve as the primarystructural member with respect to rotational and other forces caused orexperienced during the operation of engine 10, and that airfoil skin 42be designed to withstand thermal stresses and the forces related to theflow of hot gases past the outer surface of the airfoil. However,designs where airfoil skin 42 also serves as a significant structuralmember with respect to operating forces are also possible.

Second material used to form airfoil 42 comprises any suitable material,but must be capable of withstanding the operating gas temperatures of aturbine engine. Such as about 1650° C. in the case of current aircraftengines, and higher in the case of advanced designs. Airfoil supportwall 42 must be compatible with and adapted to be metallurgically bondedto the airfoil-shaped outer surface 48 of airfoil support wall 42. It isessential that airfoil skin 42 be metallurgically bonded to airfoilsupport wall 40. This metallurgical bond is preferably formed whenairfoil skin 42 is deposited onto airfoil support wall 40. This bondingmay be influences during the deposition by many parameters including themethod of deposition, the temperature of the airfoil support wall duringthe deposition, whether the deposition surface is biased relative to thedeposition source, and other parameters. Bonding may also be affected bysubsequent heat treatment or other processing. In addition, the surfacemorphology, chemistry and cleanliness of airfoil support wall 40 priorto the deposition can influence the degree to which metallurgicalbonding occurs. For example, it is known when using sputter deposition,that sputtering the surface to be deposited often improves the degree ofbonding between the substrate and subsequently deposited layers. Inaddition to forming a strong metallurgical bond between airfoil skin 42and airfoil support wall 40, it is necessary that this bond remainstable over time and at high temperatures with respect to phase changesand interdiffusion, as described herein. By compatible, it is preferredthat the metallurgical bond between these elements be thermodynamicallystable such that the strength and ductility of the bond do notdeteriorate significantly over time (e.g. up to 3 years) byinterdiffusion or other processes, even for exposures at hightemperatures on the order of 1,150° C., for Ni-base alloy airfoilsupport walls 40 and Ni-base airfoil skins 42, or higher temperatures onthe order of 1,300° C. in the case where higher temperature materialsare utilized, such as Nb-base alloys.

Where the first material of airfoil support wall 40 is an Ni-basesuperalloy containing both γ and γ' phases or a NiAl intermetallicalloy, second materials for airfoil skin 42 may comprise these samematerials. Such a combination of airfoil skin 42 and airfoil supportwall materials is preferred for applications such as where the maximumtemperatures of the operating environment similar to those of existingengines (e.g. below 1650° C.). In the case where the first material ofairfoil support wall 40 is an Nb-base alloys, second materials forairfoil skin 42 may also comprise an Nb-base alloy, including the sameNb-base alloy.

For other applications, such as applications that impose temperature,environmental or other constraints that make the use of a metal alloyairfoil skin 42 undesirable, it is preferred that airfoil skin 42comprise materials that have properties that are superior to those ofmetal alloys alone, such as composites in the general form ofintermetallic compound (I_(s))/metal alloy (M) phase composites andintermetallic compound (I_(s))/intermetallic compound (I_(M)) phasecomposites. Metal alloy M may be the same alloy as used for airfoilsupport wall 40, or a different material, depending on the requirementsof the airfoil. These composites are generally speaking similar in thatthey combine a relatively more ductile phase M or IM with a relativelyless ductile phase I_(s). The objective being to create an airfoil skin42 that gains the advantage of both materials. The less ductile materialmust have superior strength, particularly at high temperatures, and mustbe adapted to contribute this strength to the composite. It is notsufficient that the less ductile material have high strength by itself.What is important is that the less ductile material have such superiorstrength and also that it retain such strength in the composite, andafter long exposures at high temperatures. The more ductile materialmust have sufficient low temperature and high temperature ductility andmust be adapted to contribute this ductility to the less ductilematerial, and retain sufficient ductility after long exposures at hightemperatures. One factor which bears on retention of strength in acomposite is chemical compatibility of the materials of theconstituents. Thus, in order to have a successful composite, the twomaterials must be compatible. As used herein in regard to composites,the term compatible means that the materials must be capable of formingthe desired initial distribution of their phases, and of maintainingthat distribution for extended periods of time as described above at usetemperatures of 1,150° C. or more, without undergoing metallurgicalreactions that substantially impair the strength, ductility, toughness,and other important properties of the composite. Such compatibility canalso be expressed in terms of phase stability. That is, the separatephases of the composite must have a stability during operation attemperature over extended periods of time so that these phases remainseparate and distinct, retaining their separate identities andproperties and do not become a single phase or a plurality of differentphases due to interdiffusion. Compatibility can also be expressed interms of morphological stability of the interphase boundary interfacebetween the I_(S) /M or I_(S) /I_(M) composite layers. Such instabilitymay be manifested by convolutions which disrupt the continuity of eitherlayer. It is also noted that within a given airfoil skin 42, a pluralityof I_(S) /M or I_(S) /I_(M) composites may also be used, and suchcomposites are not limited to two material or two phase combinations.The use of such combinations are merely illustrative, and not exhaustiveor limiting of the potential combinations. Thus M/I_(M) /I_(S), M/I_(S1)/I_(S2) (where I_(S1) and I_(S2) are different materials) and many othercombinations are possible.

In a preferred embodiment, where airfoil skin 42 is an I_(s) /M or I_(S)/I_(M) composite, the composite will comprise a plurality of alternatinglayers of these materials. The thickness of each of the plurality oflayers may be constant or variable. It is preferred in both types ofcomposites that the volume fraction of I_(S) in the airfoil skin be inthe range of 0.3-0.7. If airfoil skin contains too little I_(S), thecomposite will not have a significant improvement in strength over I_(M)or M, if it contains too much I_(S), airfoil skin 42 will not have asignificant gain in ductility over I_(S). It is also preferred that thefirst layer that is in contact with the airfoil-shaped outer surface 48of airfoil support wall 40 be the more ductile phase, M or I_(M), tofoster metallurgical bonding with the airfoil-shaped outer surface 48 ofairfoil support wall 40. Because the materials used for M or I_(M),I_(S) and the airfoil support wall may have different coefficients ofthermal expansion, it may be preferred to form a graded airfoil skin 42,such that the portion of airfoil skin 42 closest to airfoil support wall40 has a higher volume fraction of the phase that has a thermalexpansion coefficient closest to that of the first material of airfoilsupport wall 40, to avoid exacerbating stresses induced by the expansioncoefficient mismatch that will occur during normal thermal cyclingairfoil. Similarly, it may be desirable to provide a lower volumefraction of the phase that has a higher thermal expansion coefficient inthe portion of airfoil skin 42 farthest from airfoil support wall 40.Such grading may also be used to promote certain other properties, suchas oxidation resistance, strength or other properties within specificportions of airfoil skin 42.

Since the toughening M or I_(M) phase and the strengthening I_(S) phasemust co-exist at a plurality of interfaces for long exposures at hightemperatures, stability of these interfaces will be important tomaintaining the mechanical properties of airfoil skin 42. Likewise, theinterface of airfoil skin 42 and airfoil support wall 40 must remainstable over time. The selection of the chemistries of the first andsecond materials must be made so as to achieve thermodynamic equilibriumbetween the various constituents. Some interdiffusion is necessary topromote metallurgical bonding between them, but compositions must beselected and the materials processed to avoid excessive interdiffusionleading to a loss of phase stability or substantial alteration in themechanical properties.

In composite airfoil skins 42, with respect to the various combinationsof first material for airfoil support wall 40, I_(S) and M or I_(M), thenumber of combinations is large. Small variations in the composition ofthe first material portend the possibility of large or small changes inI_(S), M or I_(M) to achieve compatibility as described above. Also,more than one acceptable combination of I_(S), M or I_(M) is likely fora given first material. In I_(S) /M composites, it is preferred in manycombinations that I_(S) and M be selected so as to comprise one or moreof the constituents of the first material, such as Ni and Cr in the caseof Ni-base alloys, so as to promote metallurgical bonding and stabilityas described herein.

For example, in the case where airfoil support wall 40 is a Ni-basesuperalloy comprising a mixture of both γ and γ' phases, I_(S) maycomprise Ni₃ [Ti, Ta, Nb, V], NiAl, Cr₃ Si, [Cr, Mo]_(X) Si, [Ta, Ti,Nb, Hf, Zr, V]C, Cr₃ C₂ and Cr₇ C₃ intermetallic compounds andintermediate phases and M may comprise a Ni-base superalloy comprising amixture of both γ and γ' phases. In Ni-base superalloys comprising amixture of both γ and γ' phases, the elements Co, Cr, Al, C and B arenearly always present as alloying constituents, as well as varyingcombinations of Ti, Ta, Nb, V, W, Mo, Re, Hf and Zr. Thus, theconstituents of the exemplary I_(S) materials described correspond toone or more materials typically found in Ni-base superalloys as may beused as first material, and thus may be adapted to achieve the phase andinterdiffusional stability described herein. As an additional example inthe case where the first material comprises NiAl intermetallic alloy, ISmay comprise Ni₃ [Ti, Ta, Nb, V], NiAl, Cr₃ Si, [Cr, Mo]_(X) Si, [Ta,Ti, Nb, Hf, Zr, V]C, Cr₃ C₂ and Cr₇ C₃ intermetallic compounds andintermediate phases and I_(M) may comprise a Ni₃ Al intermetallic alloy.Again, in NiAl intermetallic alloys, one or more of the elements Co, Cr,C and B are nearly always present as alloying constituents, as well asvarying combinations of Ti, Ta, Nb, V, W, Mo, Re, Hf and Zr. Thus, theconstituents of the exemplary I_(S) materials described correspond toone or more materials typically found in NiAl alloys as may be used asfirst material, and thus may be adapted to achieve the phase andinterdiffusional stability described herein.

As another example, in the case where airfoil support wall 40 is anNb-base alloy, including an Nb-base alloy containing at least onesecondary phase, I_(S) may comprise an Nb-containing intermetalliccompound, an Nb-containing carbide or an Nb-containing boride, and M maycomprise an Nb-base alloy. It is preferred that such I_(S) /M compositecomprises an M phase of an Nb-base alloy containing Ti such that theatomic ratio of the Ti to Nb (Ti/Nb) of the alloy is in the range of0.2-1, and an I_(S) phase comprising a group consisting of Nb-basesilicides, Cr₂ [Nb, Ti, Hf], and Nb-base aluminides, and wherein Nb,among Nb, Ti and Hf, is the primary constituent of Cr₂ [Nb, Ti, Hf]on anatomic basis. These compounds all have Nb as a common constituent, andthus may be adapted to achieve the phase and interdiffusional stabilitydescribed herein.

Airfoil skins 42 comprising composite materials have an additionaladvantage over prior art airfoils and integrally formed double-wallairfoils, related to weight, and hence engine 10 efficiency. In mostcases, intermetallic compounds and intermediate phases have densitiesthat are substantially less than metal alloys, particularly Ni-basesuperalloys. In a microlaminate configuration, the density is estimatedto be more on the order of 6.5 g/cc, rather than the 8.5-8.6 g/cc foundin common Ni-base superalloys. Also, calculations for composites, andparticularly microlaminates, indicate that they should exhibitsignificantly higher strength at a given elevated temperature thanNi-base superalloys, as shown in FIG. 4. Therefore, they can be utilizedin airfoils and engines that work at higher operating gas temperaturesand efficiencies than prior art Ni-base superalloys.

Having described the structure of an airfoil of the present invention, amethod for making these airfoils, and particularly the integral internalchannels is described.

Method of Making Airfoils

Referring to FIG. 5 and FIG. 1, a method of making double-wall airfoils16 is described comprising the steps of: forming 100 a partially-hollowairfoil support wall 40 that is attached to and extends longitudinally44 from airfoil base 18 and has airfoil-shaped outer surface 48, saidairfoil support wall 40 formed from a first material and having at leastone recessed channel 46 formed in airfoil-shaped outer surface 48;filling 110 the at least one recessed channel 46 with channel fillingmeans 112 so as to make upper surface 114 of channel filling means 112substantially continuous with airfoil-shaped outer surface 48;depositing 120 airfoil skin 42 of a second material onto theairfoil-shaped outer surface 48 of the airfoil support wall 40 such thatthe airfoil skin 42 conforms to the airfoil-shaped outer surface 48 ofthe airfoil support wall 40 and covers recessed channel 46 and channelfilling means 112, and such that airfoil skin 42 is metallurgicallybonded to airfoil-shaped outer surface 48 of airfoil support wall 40,wherein the combination of airfoil skin 42 and airfoil support wall 40form double-wall airfoil structure 50; and removing 130 channel fillingmeans 112, thereby creating integral internal channel 52 withindouble-wall airfoil structure 50.

Referring to FIG. 5, any suitable method of forming 100 airfoil supportwall 40 may be utilized, and will depend upon the first materialselected, the overall desired configuration of airfoil support wall 40(e.g. airfoil type, complexity of the internal cavities which comprisethe hollow spaces, desired channel configuration) and other factors.Airfoil 16 may be cast integrally with base 18, or otherwise attached tobase 18 as described herein. Any suitable first material may beselected, such as a metal alloy, an I_(S) /M composite or an I_(S)/I_(M) composite, as described herein. A preferred step of forming 100would comprise casting airfoil support wail 40. In the case of Ni-basesuperalloys and NiAl alloys, it is further preferred that airfoilsupport wall 40 be cast so as to form single crystal or directionallysolidified microstructures and thereby improve the high temperaturestrength and creep resistance of airfoil support wall 40. It is furtherpreferred that ail features of airfoil support wall 40, including the atleast one recessed channel 46 be formed integrally in airfoil supportwall 40 as part of the forming process, such as by casting the channelsas a feature into the airfoil-shaped outer wall 48. Such a method wouldminimize the number of steps required to form airfoil support wall 40.However, it is also possible that the at least one recessed channel beformed in a separate operation within the overall step of forming 100airfoil support wall 40. Such channel forming operations comprisemachining; such as mechanical machining with a cutting means,electro-discharge machining, water jet machining; patterning andetching, such as photo patterning and chemical etching orelectro-etching; mechanical abrasion or other suitable operations.

Referring again to FIG. 5, after forming 100 the airfoil support wall,the next step is filling 110 the at least one recessed channel 46 withchannel filling means 112. Filling 110 with channel filling means 112may be done using any suitable method, and will depend on what type ofchannel filling means 112 is utilized. Channel filling means 112 maycomprise metal alloys, ceramics, metal alloy/ceramic mixtures, polymersor other suitable means. Channel filling means 112 must first be able tobe adapted to fill the at least one recessed channel 46 so as to formupper surface 114 that is substantially continuous with airfoil-shapedouter surface 48. If channel filling means 112 is not continuous, thenit is possible to produce defects in airfoil skin 42 at the edges of therecessed channels 46 when airfoil skin 42 is subsequently deposited. Forexample, if channel filling means 112 shrinks or pulls away from therecessed channels 46, a gap would result, into which airfoil skin 42 maybe deposited. This may produce a defect in airfoil skin 42, that couldresult in an area of mechanical weakness in airfoil skin 42. For Ni-basesuperalloys, a channel filling means 112 comprising a metal alloy brazecomprising an alloy of Cu and Ni is preferred, where the alloy comprisesNi-25-35Cu, by weight. This alloy can be applied into the recessedchannels 46 so as to form an excess within the channel, such as bybrazing. The excess may be removed by machining or other suitableremoval means to form top surface 114 that is substantially continuouswith airfoil-shaped outer surface 48. Depending on the channel fillingmeans 112 characteristics at elevated temperatures, it may be desirableto control the temperature during the deposition of airfoil skin 42, soas to avoid disturbing channel filling means 112 or disrupting thecontinuity of the interface between channel filling means 112 andairfoil support wall 40, or substantial interaction between channelfilling means 112 and airfoil skin 42 or the airfoil support wall 40. Inthe case of the Ni--Cu braze described, it is thought to be desirable tomaintain the temperature during deposition at about 1200° C., or less,for as short a time as possible.

Depositing 120 of airfoil skin 42 may be done by any suitable method.Known deposition methods comprise physical vapor deposition, such asevaporation, sputtering and plasma spraying, chemical vapor deposition,including metal-organic decomposition, and plating, such aselectro-plating. Evaporation is preferred for most airfoil skinmaterials, because it generally has higher deposition rates than otherdeposition methods. In particular, electron beam evaporation ispreferred, again related to the deposition rates that are attainable. Itis most preferred to use an electron beam evaporation method developedby Applicants and described in co-pending patent application Ser. No.08/364,152 filed on Dec. 27, 1994. (Applicants Docket Number RD-22,535), which is herein incorporated by reference. This method comprisesevaporating one material through a molten pool of another material. Itis believed that a large number of materials may be evaporated using themethod described herein including, pure metals, metal alloys, ceramics,and other inorganic metallic compounds. However, at the present time itis particularly preferable to use the method for evaporative depositionof multi-constituent metal alloys.

This method enables the making an evaporated deposit of a material, bythe steps of: selecting a first material and a second material having acomposition comprising a plurality of elements, wherein the selectionensures that the first material is adapted to transport the plurality ofelements of the second material through the first material when they arein touching contact with one another in a molten state and that theplurality of elements of the molten second material in molten firstmaterial are preferentially evaporated with respect to the firstmaterial; placing a quantity of the first material over a quantity of asecond material in a crucible means contained within a housing meansthat is adapted to be evacuated so that the first material at leastpartially covers the second material; evacuating the housing means;supplying heat to the first material sufficient to create a molten zonewithin and through this material such that the molten zone of the firstmaterial is in touching contact with the second material and therebycreates a molten zone within the second material, wherein the pluralityof elements of the second material are transported through the moltenzone rich in the first material to a top surface where they arepreferentially evaporated with respect to the first material therebyforming a vapor stream; and collecting a condensate having a thicknessfrom the vapor stream on a collection means, wherein the composition ofthe condensate closely resembles the composition of the second materialthroughout the thickness of the condensate.

With respect to many high temperature materials, this method may bedescribed as a method of making an evaporated deposit of a material,comprising the steps of: selecting a first material having a firstmelting point, and a second material having a composition comprising aplurality of elements and having a second melting point, wherein theselection ensures that the first melting point is greater than thesecond melting point, and that the first material is adapted totransport the plurality of elements of the second material through thefirst material when they are in touching contact with one another in amolten state, and that the plurality of elements of the molten secondmaterial in molten first material are preferentially evaporated withrespect to the first material; placing a quantity of the first materialover a quantity of a second material in a crucible means containedwithin a housing means that is adapted to be evacuated so that the firstmaterial at least partially covers the second material; evacuating thehousing means; supplying heat to the first material sufficient to createa molten zone within and through this material such that the molten zoneof the first material is in touching contact with the second materialand thereby creates a molten zone within the second material, whereinthe plurality of elements of the second material are transported throughthe molten zone of the first material to a top surface where they arepreferentially evaporated with respect to the first material therebyforming a vapor stream; and collecting a condensate having a thicknessfrom the vapor stream on a collection means, wherein the composition ofthe condensate closely resembles the composition of the second materialthroughout the thickness of the condensate.

It is also possible to adapt this evaporation method such that it may beused to make an evaporated deposit of a material by the steps of:selecting a first material and a second material having a compositioncomprising a plurality of elements, wherein the selection ensures thatthe first material is adapted to transport the plurality of elements ofthe second material through the first material when they are in touchingcontact with one another in a molten state and that at least one of theplurality of elements of the molten second material in molten fastmaterial is preferentially evaporated with respect to the firstmaterial; placing a quantity of the first material over a quantity of asecond material in a crucible means contained within a housing meansthat is adapted to be evacuated so that the first material at leastpartially covers the second material; evacuating the housing means;supplying heat to the first material sufficient to create a molten zonewithin and through this material such that the molten zone of the firstmaterial is in touching contact with the second material and therebycreates a molten zone within the second material, wherein the pluralityof elements of the second material are transported through the moltenzone rich in the first material to a top surface where they arepreferentially evaporated with respect to the first material therebyforming a vapor stream; and collecting a condensate having a thicknessfrom the vapor stream on a collection means, wherein the composition ofthe condensate closely resembles the composition of the second materialthroughout the thickness of the condensate, except that the condensatealso contains controlled quantifies of the first material throughout thethickness.

In a preferred embodiment, the first material is a refractory metal,such as W, Re, Os, Ta, Mo, Nb, Ir, Ru, or Hf, or alloys of these metals,and the second material is a multi-constituent metal alloy, such as aNi-base, Co-base or Fe-base alloy.

This evaporation method is particularly advantageous in that it permitsthe evaporation and subsequent collection as a condensate of secondmaterials that have a composition closely resembling the startingmaterial from which they were evaporated and exclusive of the elementsin the first material. One of the significant problems in the use ofevaporative techniques generally, is that without elaborate controls,the condensate of multi-constituent material usually does not have acomposition that closely resembles the composition of the startingmaterial. This is due to the fact that at a particular temperature thevarious elements of a material comprising a plurality of elementsevaporate at differing rates, which are related to the vapor pressuresof those elements at the evaporation temperature. However, thisevaporation method permits the collection of a condensate having acomposition that closely resembles the composition of the startingmaterial, without the need for elaborate compositionally orientedcontrols over the deposition process.

Another significant advantage in the use of this technique is that itpermits the deposition of the compositionally controlled condensates atrates that are very high for an evaporative process, on the order of atleast 0.5 mil/minute, and often 1 mil/minute or more.

The combination of these advantages provides a method for theevaporative deposition of high temperature alloys and other materialsfor uses in structural applications, such as airfoil skin 42.

Yet another advantage of this evaporation technique is that thepropensity of the molten pool from which the material is evaporated tosplatter during the course of the deposition is greatly reduced.

The step of removing 130 channel filling means 112 may be done using anysuitable method, such as melting/extraction, chemical etching, pyrolysisor other formation of volatile reaction products, or other methods. Inthe case of metal alloy brazes, the braze may be either removed bymelting/extraction or etching.

What is claimed is:
 1. An airfoil having an outer double-wall,comprising;a partially hollow airfoil support wall that is attached toand extends longitudinally from an airfoil base and has anairfoil-shaped outer surface, the airfoil support wall formed from afirst material and having at least one longitudinally-extending recessedchannel formed in the airfoil-shaped outer surface; and an airfoil skinmade from a second material which conforms to, covers and ismetallurgically bonded to the airfoil-shaped outer surface of theairfoil support wall and covers the at least onelongitudinally-extending recessed channel, wherein the combination ofthe airfoil skin and the airfoil support wall form a double-wall airfoilstructure and the covered, recessed channel forms at least onelongitudinally-extending integral internal channel located within thedouble-wall, and wherein the second material comprises an I_(s) /M orI_(s) /I_(M) composite in the form of a micro-laminate having aplurality of alternating layers of I_(s) and M or I_(M).
 2. The airfoilof claim 1 wherein the first material comprises a metal alloy, an I_(s)/M composite or an I_(s) /I_(M) composite.
 3. The airfoil of claim 1wherein the first material is a metal alloy having a single crystal ordirectionally solidified microstructure.
 4. The airfoil of claim 1wherein the second material is an I_(s) /M composite.
 5. The airfoil ofclaim 4 wherein a total combined thickness of the I_(s) layers isI_(SSUM) and a total combined thickness of the M layers is M_(SUM), andthe ratio I_(SSUM) /I_(SSUM) +M_(SUM) is in the range of about 0.3-0.7.6. The airfoil of claim 5 wherein a first layer of the second materialforming said airfoil skin is an M layer and is metallurgically bonded tothe airfoil-shaped outer surface of the airfoil support wall.
 7. Theairfoil skin of claim 4 wherein the arrangement of the plurality ofalternating layers of I_(s) and M is graded to provide that the airfoilskin contains a higher volume fraction of the layers of either I_(s) orM, whichever has a coefficient of thermal expansion that most closelymatches a coefficient of thermal expansion of the first material, in aportion of the airfoil skin adjacent to the airfoil support wall.
 8. Theairfoil of claim 1 wherein the second material is an I_(s) /I_(M)composite.
 9. The airfoil of claim 8 wherein a total combined thicknessof the I_(s) layers is I_(SSUM) and a total combined thickness of theI_(M) layers Is I_(MSUM), and the ratio I_(SSUM) /I_(SSUM) +I_(MSUM) isin the range of about 0.3-0.7.
 10. The airfoil of claim 9 wherein afirst layer of the second material forming said airfoil skin is an I_(M)layer and is metallurgically bonded to the airfoil-shaped outer surfaceof the airfoil support wall.
 11. The airfoil of claim 1 wherein thefirst material is from a group consisting of Ni-base superalloyscontaining γ and γ' phases and NiAl intermetallic alloys, and the secondmaterial is from a group consisting of Ni-base superalloys containing γand γ' phases, NiAl intermetallic alloys, I_(s) /M composites and I_(s)/I_(M) composites.
 12. The airfoil of claim 11 wherein the secondmaterial is an I_(s) /M composite comprising an I_(s) from a groupconsisting of Ni₃ [Ti,Ta,Nb,V], NiAl, Cr₃ Si, [Cr,Mo]_(X) Si,[Ta,Ti,Nb,Hf,Zr,V]C, Cr₃ C₂ and Cr₇ C₃ intermetallic compounds andintermediate phases and an M of a Ni-base superalloy comprising amixture of γ and γ' phases.
 13. The airfoil of claim 11 wherein thesecond material is an I_(s) /I_(M) composite comprising an I_(s) from agroup consisting of Ni₃ [Ti,Ta,Nb,V], NiAl, Cr₃ Si, [Cr,Mo]_(X) Si,[Ta,Ti,Nb,Hf,Zr,V]C, Cr₃ C₂ and Cr₇ C₃ intermetallic compounds andintermediate phases and an I_(M) of a Ni₃ Al intermetallic compound. 14.The airfoil of claim 1 wherein the first material is an Nb-base alloyand the second material is an I_(s) /M composite comprising an I_(s)from a group consisting of Nb-containing intermetallic compounds,Nb-containing carbides and Nb-containing borides and an M of an Nb-basealloy.
 15. The airfoil of claim 14 wherein the first material is anNb-base alloy containing Ti such that the atomic ratio of the Ti to Nb(Ti/Nb) of the alloy is in the range of about 0.5-1.
 16. The airfoil ofclaim 14 wherein the I_(s) /M composite comprises an M of an Nb-basealloy containing Ti such that the atomic ratio of the Ti to Nb (Ti/Nb)of the alloy is in the range of about 0.2-1, and an I_(s) from a groupconsisting of Nb-base silicides, Cr₂ [Nb, Ti, Hf], and Nb-basealuminides, and wherein Nb is the primary constituent of Cr₂ [Nb, Ti,Hf] on an atomic basis.
 17. The airfoil of claim 1 wherein the at leastone channel comprises a plurality of channels.
 18. The airfoil of claim17 wherein the plurality of channels covers at least 50 percent of theouter surface of the airfoil support wall.
 19. The airfoil of claim 1wherein the at least one channel has a serpentine pattern.
 20. Theairfoil of claim 1 wherein said airfoil skin has a thickness in therange of about 0.010-0.040 inches.